Geared architecture for high speed and small volume fan drive turbine

ABSTRACT

A turbofan engine includes a propulsor section that has a propulsor shaft in driving engagement with a propulsor. An epicyclic gear system has a gear mesh lateral stiffness and a gear mesh transverse stiffness. A gear system input defines a gear system input lateral stiffness and a gear system input transverse stiffness. The gear system input lateral stiffness is less than 5% of the gear mesh lateral stiffness. A first turbine section rotates at a first speed, and a second turbine rotates at a second speed that is faster than the first speed. A first performance quantity is defined as the product of the first speed squared and the first area of the first turbine, a second performance quantity is defined as the product of the second speed squared and the second area of the second turbine, and a performance quantity ratio is between 0.5 and 1.5.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/342,100, filed Jun. 8, 2021, which is a continuation of U.S. patentapplication Ser. No. 15/881,240, filed Jan. 26, 2018, which is acontinuation of U.S. patent application Ser. No. 15/485,512, filed Apr.12, 2017 which is a continuation of U.S. patent application Ser. No.13/908,177, filed Jun. 3, 2013, now U.S. Pat. No. 9,631,558 granted Apr.25, 2017, which is a continuation-in-part of U.S. patent applicationSer. No. 13/623,309, filed Sep. 20, 2012, now U.S. Pat. No. 9,133,729,granted Sep. 15, 2015, which is a continuation-in-part of U.S. patentapplication Ser. No. 13/342,508, filed Jan. 3, 2012, now U.S. Pat. No.8,297,916, granted Oct. 30, 2012, and which claims priority to U.S.Provisional Patent Application No. 61/494,453, filed Jun. 8, 2011.

BACKGROUND

The present disclosure relates to a gas turbine engine, and moreparticularly to a flexible support structure for a geared architecturetherefor.

Epicyclic gearboxes with planetary or star gear trains may be used ingas turbine engines for their compact designs and efficient high gearreduction capabilities. Planetary and star gear trains generally includethree gear train elements: a central sun gear, an outer ring gear withinternal gear teeth, and a plurality of planet gears supported by aplanet carrier between and in meshed engagement with both the sun gearand the ring gear. The gear train elements share a common longitudinalcentral axis, about which at least two rotate. An advantage of epicyclicgear trains is that a rotary input can be connected to any one of thethree elements. One of the other two elements is then held stationarywith respect to the other two to permit the third to serve as an output.

In gas turbine engine applications, where a speed reduction transmissionis required, the central sun gear generally receives rotary input fromthe power plant, the outer ring gear is generally held stationary andthe planet gear carrier rotates in the same direction as the sun gear toprovide torque output at a reduced rotational speed. In star geartrains, the planet carrier is held stationary and the output shaft isdriven by the ring gear in a direction opposite that of the sun gear.

During flight, light weight structural cases deflect with aero andmaneuver loads causing significant amounts of transverse deflectioncommonly known as backbone bending of the engine. This deflection maycause the individual sun or planet gear's axis of rotation to loseparallelism with the central axis. This deflection may result in somemisalignment at gear train journal bearings and at the gear teeth mesh,which may lead to efficiency losses from the misalignment and potentialreduced life from increases in the concentrated stresses.

Further, with the geared architecture as set forth above, the torque andspeed of the input into the gear is quite high.

SUMMARY

A turbofan engine according to an example of the present disclosureincludes a propulsor section that has a propulsor shaft in drivingengagement with a propulsor having blades extending from a propulsor huband that is rotatable about a central longitudinal axis of the turbofanengine. An epicyclic gear system is in driving engagement with thepropulsor shaft and has a gear mesh lateral stiffness and a gear meshtransverse stiffness. A gear system input to the epicyclic gear systemdefines a gear system input lateral stiffness and a gear system inputtransverse stiffness. The gear system input lateral stiffness is lessthan 5% of the gear mesh lateral stiffness. A first turbine section isin driving engagement with the gear system input and is rotatable aboutthe central longitudinal axis. The first turbine section includes aninlet, an outlet, and a first exit area at a first exit point at theoutlet and rotates at a first speed. A second turbine section has asecond exit area at a second exit point and rotates at a second speed,which is faster than the first speed. The first and second speeds areredline speeds. A first performance quantity is defined as the productof the first speed squared and the first area. A second performancequantity is defined as the product of the second speed squared and thesecond area. A performance quantity ratio of the first performancequantity to the second performance quantity is between 0.5 and 1.5.

In a further embodiment of any of the foregoing embodiments, theepicyclic gear system includes a ring gear having a ring gear lateralstiffness and a ring gear transverse stiffness, and the ring gearlateral stiffness is less than 12% of the gear mesh lateral stiffness.

The turbofan engine as set forth in claim 2, wherein the first turbinehas between three and six stages and the second turbine has two stages.

In a further embodiment of any of the foregoing embodiments, the secondspeed is more than twice the first speed.

In a further embodiment of any of the foregoing embodiments, there is aflexible support to support the epicyclic gear system, the flexiblesupport has a flexible support lateral stiffness and a flexible supporttransverse stiffness, a frame which supports the fan shaft has a framelateral stiffness and a frame transverse stiffness, and at least one ofthe flexible support lateral stiffness and the flexible supporttransverse stiffness are less than 11% of a respective one of the framelateral stiffness and the frame transverse stiffness.

In a further embodiment of any of the foregoing embodiments, the gearsystem input connects a sun gear of the epicyclic gear system to bedriven by the first turbine section, and the gear system input lateralstiffness and the gear system input transverse stiffness is less than11% of a respective one of the frame lateral stiffness and the frametransverse stiffness.

In a further embodiment of any of the foregoing embodiments, theperformance quantity ratio is greater than or equal to 0.5 and less thanor equal to 1.075.

In a further embodiment of any of the foregoing embodiments, the gearsystem input transverse stiffness is less than 5% of the gear meshtransverse stiffness.

In a further embodiment of any of the foregoing embodiments, the firstturbine has between three and six stages and the second turbine has twostages.

In a further embodiment of any of the foregoing embodiments, there is aflexible support to support the epicyclic gear system, the flexiblesupport has a flexible support lateral stiffness and a flexible supporttransverse stiffness, and at least one of the flexible support lateralstiffness and the flexible support transverse stiffness is less than 8%of a respective one of the gear mesh lateral stiffness and the gear meshtransverse stiffness.

In a further embodiment of any of the foregoing embodiments, there is aflexible support to support the epicyclic gear system, the flexiblesupport having a flexible support lateral stiffness and a flexiblesupport transverse stiffness, and the flexible support lateral stiffnessand the flexible support transverse stiffness being less than 8% of arespective one of the gear mesh lateral stiffness and the gear meshtransverse stiffness.

In a further embodiment of any of the foregoing embodiments, thepropulsor section is a fan section, the propulsor is a fan, thepropulsor hub is a fan hub, and an outer housing surrounds the fan todefine a bypass duct.

A turbofan engine according to an example of the present disclosureincludes a propulsor section that has a propulsor shaft that is indriving engagement with a propulsor having blades extending from apropulsor hub and are rotatable about a central longitudinal axis of theturbofan engine. An epicyclic gear system drives engagement with thepropulsor shaft and having a gear mesh lateral stiffness and a gear meshtransverse stiffness. A gear system input to the epicyclic gear systemdefines a gear system input lateral stiffness and a gear system inputtransverse stiffness. The gear system input transverse stiffness is lessthan 5% of the gear mesh transverse stiffness. A first turbine sectionis in driving engagement with the gear system input and is rotatableabout the central longitudinal axis. the first turbine section includesan inlet, an outlet, and a first exit area at a first exit point at theoutlet and rotates at a first speed. A second turbine section has asecond exit area at a second exit point and rotates at a second speed,which is faster than the first speed. The first and second speeds areredline speeds. A first performance quantity is defined as the productof the first speed squared and the first area. A second performancequantity is defined as the product of the second speed squared and thesecond area. A performance quantity ratio of the first performancequantity to the second performance quantity is between 0.5 and 1.5.

In a further embodiment of any of the foregoing embodiments, theepicyclic gear system includes a ring gear has a ring gear lateralstiffness and a ring gear transverse stiffness and the ring gear lateralstiffness is less than 12% of the gear mesh lateral stiffness.

In a further embodiment of any of the foregoing embodiments, there is aflexible support to support the epicyclic gear system, the flexiblesupport has a flexible support lateral stiffness and a flexible supporttransverse stiffness, and at least one of the flexible support lateralstiffness and the flexible support transverse stiffness is less than 8%of a respective one of the gear mesh lateral stiffness and the gear meshtransverse stiffness.

In a further embodiment of any of the foregoing embodiments, the firstturbine has between three and six stages and the second turbine has twostages.

In a further embodiment of any of the foregoing embodiments, gear systeminput lateral stiffness is less than 5% of the gear mesh lateralstiffness.

In a further embodiment of any of the foregoing embodiments, at leastone of the flexible support lateral stiffness and the flexible supporttransverse stiffness is less than 11% of a respective one of a framelateral stiffness and a frame transverse stiffness of a frame whichsupports the fan shaft.

In a further embodiment of any of the foregoing embodiments, the secondspeed is more than twice the first speed.

In a further embodiment of any of the foregoing embodiments, at leastone of the flexible support lateral stiffness and the flexible supporttransverse stiffness is less than 11% of a respective one of a framelateral stiffness and a frame transverse stiffness of a frame whichsupports the fan shaft.

In a further embodiment of any of the foregoing embodiments, a framewhich supports the fan shaft defines a frame lateral stiffness and aframe transverse stiffness, and the gear system input lateral stiffnessand the gear system input transverse stiffness are less than 11% of arespective one of the frame lateral stiffness and the frame transversestiffness.

In a further embodiment of any of the foregoing embodiments, the gearsystem input transverse stiffness is less than 5% of the gear meshtransverse stiffness.

In a further embodiment of any of the foregoing embodiments, the firstturbine section has between three and six stages and the second turbinesection has two stages.

In a further embodiment of any of the foregoing embodiments, theepicyclic gear system includes a ring gear having a ring gear lateralstiffness and a ring gear transverse stiffness and the ring gear lateralstiffness is less than 12% of the gear mesh lateral stiffness.

In a further embodiment of any of the foregoing embodiments, the ringgear transverse stiffness is less than 12% of the gear mesh transversestiffness.

In a further embodiment of any of the foregoing embodiments, there is aflexible support to support the epicyclic gear system, the flexiblesupport has a flexible support lateral stiffness and a flexible supporttransverse stiffness, and at least one of the flexible support lateralstiffness and the flexible support transverse stiffness is less than 8%of a respective one of the gear mesh lateral stiffness and the gear meshtransverse stiffness.

In a further embodiment of any of the foregoing embodiments, at leastone of the flexible support lateral stiffness and the flexible supporttransverse stiffness is less than 11% of a respective one of a framelateral stiffness and a frame transverse stiffness of a frame whichsupports the fan shaft.

In a further embodiment of any of the foregoing embodiments, a framewhich supports the fan shaft defines a frame lateral stiffness and aframe transverse stiffness and the gear system input lateral stiffnessand the gear system input transverse stiffness are less than 11% of arespective one of the frame lateral stiffness and the frame transversestiffness.

In a further embodiment of any of the foregoing embodiments, theperformance quantity ratio is above or equal to 0.8 and less than orequal to 1.075.

In a further embodiment of any of the foregoing embodiments, thepropulsor section is a fan section, the propulsor is a fan, thepropulsor hub is a fan hub, and an outer housing surrounds the fan todefine a bypass duct.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1A is a schematic cross-section of a gas turbine engine;

FIG. 1B shows a feature of the FIG. 1A engine.

FIG. 1C shows another feature.

FIG. 1D shows yet another feature.

FIG. 2 is an enlarged cross-section of a section of the gas turbineengine which illustrates a fan drive gear system (FDGS);

FIG. 3 is a schematic view of a flex mount arrangement for onenon-limiting embodiment of the FDGS;

FIG. 4 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of the FDGS;

FIG. 5 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a star system FDGS; and

FIG. 6 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a planetary system FDGS.

FIG. 7 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a star system FDGS; and

FIG. 8 is a schematic view of a flex mount arrangement for anothernon-limiting embodiment of a planetary system FDGS.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. A mid-turbine frame 57 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 57 further supports bearing systems 38in the turbine section 28. The inner shaft 40 and the outer shaft 50 areconcentric and rotate via bearing systems 38 about the engine centrallongitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram° R)/(518.7°R)]0.5. The “Low corrected fan tip speed” as disclosed herein accordingto one non-limiting embodiment is less than about 1150 ft/second.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion of the airflow passing therethrough.

The amount of thrust that can be produced by a particular turbinesection compared to how compact the turbine section is, is referred toas the power density, or the force density, of the turbine section, andis derived by the flat-rated Sea Level Take-Off (SLTO) thrust divided bythe volume of the entire turbine section. The example volume isdetermined from an inlet of the high pressure turbine 54 to an exit ofthe low pressure turbine 46. In order to increase the power density ofthe turbine section 28, each of the low pressure and high pressureturbines 46, 54 is made more compact. That is, the high pressure turbine54 and the low pressure turbine 46 are made with a shorter axial length,and the spacing between each of the turbines 46, 54 is decreased,thereby decreasing the volume of the turbine section 28.

The power density in the disclosed gas turbine engine 20 including thegear driven fan section 22 is greater than those provided in prior artgas turbine engine including a gear driven fan. Eight disclosedexemplary engines, which incorporate turbine sections and fan sectionsdriven through a reduction gear system and architectures as set forth inthis application, are described in Table I as follows:

TABLE 1 Thrust SLTO Turbine section volume Thrust/turbine section Engine(lbf) from the Inlet volume (lbf/in3) 1 17,000 3,859 4.4 2 23,300 5,3304.37 3 29,500 6,745 4.37 4 33,000 6,745 4.84 5 96,500 31,086 3.1 696,500 62,172 1.55 7 96,500 46,629 2.07 8 37,098 6,745 5.50

In some embodiments, the power density is greater than or equal to about1.5 lbf/in3. In further embodiments, the power density is greater thanor equal to about 2.0 lbf/in3. In further embodiments, the power densityis greater than or equal to about 3.0 lbf/in3. In further embodiments,the power density is greater than or equal to about 4.0 lbf in3. Infurther embodiments, the power density is less than or equal to about5.5 lbf/in3.

Engines made with the disclosed gear driven fan architecture, andincluding turbine sections as set forth in this application, providevery high efficiency operation, and increased fuel efficiency.

Referring to FIG. 1B, with continued reference to FIG. 1A, relativerotations between components of example disclosed engine architecture100 are schematically shown. In the example engine architecture 100, thefan 42 is connected, through the gearbox 48, to the low spool 30 towhich the low pressure compressor 44 and the low pressure turbine 46 areconnected. The high pressure compressor 52 and the high pressure turbine54 are connected to a common shaft forming the high spool 32. The highspool 32 rotates opposite the direction of rotation of the fan 42(illustrated in FIG. 1B as the “+” direction.) The low spool 30 rotatesin the same direction as the fan 42 (illustrated in FIG. 1B as the “−”direction.) The high pressure turbine 54 and the low pressure turbine46, along with the mid-turbine frame 57 together forms the turbinesection 28 of the gas turbine engine 20. Other relative rotationdirections between the two spools and the fan come within the scope ofthis disclosure.

One disclosed example speed change device 48 has a gear reduction ratioexceeding 2.3:1, meaning that the low pressure turbine 46 turns at least2.3 times faster than the fan 42. An example disclosed speed changedevice is an epicyclical gearbox of a planet type, where the input is tothe center “sun” gear 260. Planet gears 262 (only one shown) around thesun gear 260 rotate and are spaced apart by a carrier 264 that rotatesin a direction common to the sun gear 260. A ring gear 266, which isnon-rotatably fixed to the engine static casing 36 (shown in FIG. 1 ),contains the entire gear assembly. The fan 42 is attached to and drivenby the carrier 264 such that the direction of rotation of the fan 42 isthe same as the direction of rotation of the carrier 264 that, in turn,is the same as the direction of rotation of the input sun gear 260.Accordingly, the low pressure compressor 44 and the low pressure turbine46 counter-rotate relative to the high pressure compressor 52 and thehigh pressure turbine 54.

Counter rotating the low pressure compressor 44 and the low pressureturbine 46 relative to the high pressure compressor 52 and the highpressure turbine 54 provides certain efficient aerodynamic conditions inthe turbine section 28 as the generated high speed exhaust gas flowmoves from the high pressure turbine 54 to the low pressure turbine 46.Moreover, the mid-turbine frame 57 contributes to the overallcompactness of the turbine section 28. Further, the airfoil 59 of themid-turbine frame 57 surrounds internal bearing support structures andoil tubes that are cooled. The airfoil 59 also directs flow around theinternal bearing support structures and oil tubes for streamlining thehigh speed exhaust gas flow. Additionally, the airfoil 59 directs flowexiting the high pressure turbine 54 to a proper angle desired topromote increased efficiency of the low pressure turbine 46.

Flow exiting the high pressure turbine 54 has a significant component oftangential swirl. The flow direction exiting the high pressure turbine54 is set almost ideally for the blades in a first stage of the lowpressure turbine 46 for a wide range of engine power settings. Thus, theaerodynamic turning function of the mid turbine frame 57 can beefficiently achieved without dramatic additional alignment of airflowexiting the high pressure turbine 54.

Referring to FIG. 1C, the example turbine section 28 volume isschematically shown and includes first, second and third stages 46A, 46Band 46C. Each of the stages 46A, 46B and 46C includes a correspondingplurality of blades 212 and vanes 214. The example turbine sectionfurther includes an example air-turning vane 220 between the low andhigh turbines 54, 46 that has a modest camber to provide a small degreeof redirection and achieve a desired flow angle relative to blades 212of the first stage 46 a of the low pressure turbine 46. The disclosedvane 220 could not efficiently perform the desired airflow function ifthe low and high pressure turbines 54, 46 rotated in a common direction.

The example mid-turbine frame 57 includes multiple air turning vanes 220in a row that direct air flow exiting the high pressure turbine 54 andensure that air is flowing in the proper direction and with the properamount of swirl. Because the disclosed turbine section 28 is morecompact than previously utilized turbine sections, air has less distanceto travel between exiting the mid-turbine frame 57 and entering the lowpressure turbine 46. The smaller axial travel distance results in adecrease in the amount of swirl lost by the airflow during thetransition from the mid-turbine frame 57 to the low pressure turbine 46,and allows the vanes 220 of the mid-turbine frame 57 to function asinlet guide vanes of the low pressure turbine 46. The mid-turbine frame57 also includes a strut 221 providing structural support to both themid-turbine frame 57 and to the engine housing. In one example, themid-turbine frame 57 is much more compact by encasing the strut 221within the vane 220, thereby decreasing the length of the mid-turbineframe 57.

At a given fan tip speed and thrust level provided by a given fan size,the inclusion of the speed change device 48 (shown in FIGS. 1A and 1B)provides a gear reduction ratio, and thus the speed of the low pressureturbine 46 and low pressure compressor 44 components may be increased.More specifically, for a given fan diameter and fan tip speed, increasesin gear ratios provide for a faster turning turbine that, in turn,provides for an increasingly compact turbine and increased thrust tovolume ratios of the turbine section 28. By increasing the gearreduction ratio, the speed at which the low pressure compressor 44 andthe low pressure turbine 46 turn, relative to the speed of the fan 42,is increased.

Increases in rotational speeds of the gas turbine engine 20 componentsincreases overall efficiency, thereby providing for reductions in thediameter and the number of stages of the low pressure compressor 44 andthe low pressure turbine 46 that would otherwise be required to maintaindesired flow characteristics of the air flowing through the core flowpath C. The axial length of each of the low pressure compressor 44 andthe low pressure turbine 46 can therefore be further reduced due toefficiencies gained from increased speed provided by an increased gearratio. Moreover, the reduction in the diameter and the stage count ofthe turbine section 28 increases the compactness and provides for anoverall decrease in required axial length of the example gas turbineengine 20.

In order to further improve the thrust density of the gas turbine engine20, the example turbine section 28 (including the high pressure turbine54, the mid-turbine frame 57, and the low pressure turbine 46) is mademore compact than traditional turbine engine designs, thereby decreasingthe length of the turbine section 28 and the overall length of the gasturbine engine 20.

In order to make the example low pressure turbine 46 compact, make thediameter of the low pressure turbine 46 more compatible with the highpressure turbine 54, and thereby make the air-turning vane 220 of themid-turbine frame 57 practical, stronger materials in the initial stagesof the low pressure turbine 46 may be required. The speeds andcentrifugal pull generated at the compact diameter of the low pressureturbine 46 pose a challenge to materials used in prior art low pressureturbines.

Examples of materials and processes within the contemplation of thisdisclosure for the air-turning vane 220, the low pressure turbine blades212, and the vanes 214 include materials with directionally solidifiedgrains to provided added strength in a span-wise direction. An examplemethod for creating a vane 220, 214 or turbine blade 212 havingdirectionally solidified grains can be found in U.S. application Ser.No. 13/290,667, and U.S. Pat. Nos. 7,338,259 and 7,871,247, each ofwhich is incorporated by reference. A further, engine embodimentutilizes a cast, hollow blade 212 or vane 214 with cooling airintroduced at the leading edge of the blade/vane and a trailing edgedischarge of the cooling air. Another embodiment uses an internallycooled blade 212 or vane 214 with film cooling holes. An additionalengine embodiment utilizes an aluminum lithium material for constructionof a portion of the low pressure turbine 46. The example low pressureturbine 46 may also be constructed utilizing at a powdered metal disc orrotor.

Additionally, one or more rows of turbine blades 212 of the low pressureturbine 46 can be constructed using a single crystal blade material.Single crystal constructions oxidize at higher temperatures as comparedto non-single crystal constructions and thus can withstand highertemperature airflow. Higher temperature capability of the turbine blades212 provide for a more efficient low pressure turbine 46 that may befurther reduced in size.

While the illustrated low pressure turbine 46 includes three turbinestages 46 a, 46 b, and 46 c, the low pressure turbine 46 can be modifiedto include up to six turbine stages. Increasing the number of lowpressure turbine stages 46 a, 46 b, 46 c at constant thrust slightlyreduces the thrust density of the turbine section 28 but also increasespower available to drive the low pressure compressor and the fan section22.

Further, the example turbine blades may be internally cooled to allowthe material to retain a desired strength at higher temperatures andthereby perform as desired in view of the increased centrifugal forcegenerated by the compact configuration while also withstanding thehigher temperatures created by adding low pressure compressor 44 stagesand increasing fan tip diameter.

Each of the disclosed embodiments enables the low pressure turbine 46 tobe more compact and efficient, while also improving radial alignment tothe high pressure turbine 54. Improved radial alignment between the lowand high pressure turbines 54, 46 increases efficiencies that can offsetany increases in manufacturing costs incurred by including the airturning vane 220 of the mid-turbine frame 57.

In light of the foregoing embodiments, the overall size of the turbinesection 28 has been greatly reduced, thereby enhancing the engine'spower density. Further, as a result of the improvement in power density,the engine's overall propulsive efficiency has been improved.

An exit area 401 is shown, in FIG. 1D and FIG. 1A, at the exit locationfor the high pressure turbine section 54. An exit area for the lowpressure turbine section is defined at exit 401 for the low pressureturbine section. As shown in FIG. 1D, the turbine engine 20 may becounter-rotating. This means that the low pressure turbine section 46and low pressure compressor section 44 rotate in one direction, whilethe high pressure spool 32, including high pressure turbine section 54and high pressure compressor section 52 rotate in an opposed direction.The gear reduction 48, which may be, for example, an epicyclictransmission (e.g., with a sun, ring, and star gears), is selected suchthat the fan 42 rotates in the same direction as the high spool 32. Withthis arrangement, and with the other structure as set forth above,including the various quantities and operational ranges, a very highspeed can be provided to the low pressure spool. Low pressure turbinesection and high pressure turbine section operation are often evaluatedlooking at a performance quantity which is the exit area for the turbinesection multiplied by its respective speed squared. This performancequantity (“PQ”) is defined as:

PQlpt=(Alpt×Vlpt2)  Equation 1:

PQhpt=(Ahpt×Vhpt2)  Equation 2:

where Alpt is the area of the low pressure turbine section at the exitthereof (e.g., at 401), where Vlpt is the speed of the low pressureturbine section, where Ahpt is the area of the high pressure turbinesection at the exit thereof (e.g., at 400), and where Vhpt is the speedof the low pressure turbine section.

Thus, a ratio of the performance quantity for the low pressure turbinesection compared to the performance quantify for the high pressureturbine section is:

(Alptp33 Vlpt2)/(Ahpt×Vhpt2)=PQlpt/PQhpt  Equation 3:

In one turbine embodiment made according to the above design, the areasof the low and high pressure turbine sections are 557.9 in2 and 90.67in2, respectively. Further, the speeds of the low and high pressureturbine sections are 10179 rpm and 24346 rpm, respectively. Thus, usingEquations 1 and 2 above, the performance quantities for the low and highpressure turbine sections are:

PQlpt=(Alpt×Vlpt2)=(557.9 in2)(10179 rpm)2=57805157673.9 in2rpm2  Equation 1:

PQhpt=(Ahpt×Vhpt2)=(90.67 in2)(24346 rpm)2=53742622009.72 in2rpm2  Equation 2:

and using Equation 3 above, the ratio for the low pressure turbinesection to the high pressure turbine section is:

Ratio=PQlpt/PQhpt=57805157673.9 in2 rpm2/53742622009.72 in2 rpm2=1.075

In another embodiment, the ratio was about 0.5 and in another embodimentthe ratio was about 1.5. With PQlpt/PQhpt ratios in the 0.5 to 1.5range, a very efficient overall gas turbine engine is achieved. Morenarrowly, PQlpt/PQhpt ratios of above or equal to about 0.8 are moreefficient. Even more narrowly, PQlpt/PQhpt ratios above or equal to 1.0are even more efficient. As a result of these PQlpt/PQhpt ratios, inparticular, the turbine section can be made much smaller than in theprior art, both in diameter and axial length. In addition, theefficiency of the overall engine is greatly increased.

The low pressure compressor section is also improved with thisarrangement, and behaves more like a high pressure compressor sectionthan a traditional low pressure compressor section. It is more efficientthan the prior art, and can provide more work in fewer stages. The lowpressure compressor section may be made smaller in radius and shorter inlength while contributing more toward achieving the overall pressureratio design target of the engine.

A worker of ordinary skill in the art, being apprised of the disclosureabove, would recognize that high torque and high speed will be presentedby the low speed spool 30 into the gear architecture 48. Thus, aflexible mount arrangement becomes important.

With reference to FIG. 2 , the geared architecture 48 generally includesa fan drive gear system (FDGS) 60 driven by the low speed spool 30(illustrated schematically) through an input coupling 62. The inputcoupling 62 both transfers torque from the low speed spool 30 to thegeared architecture 48 and facilitates the segregation of vibrations andother transients therebetween. In the disclosed non-limiting embodiment,the FDGS 60 may include an epicyclic gear system which may be, forexample, a star system or a planet system.

The input coupling 62 may include an interface spline 64 joined, by agear spline 66, to a sun gear 68 of the FDGS 60. The sun gear 68 is inmeshed engagement with multiple planet gears 70, of which theillustrated planet gear 70 is representative. Each planet gear 70 isrotatably mounted in a planet carrier 72 by a respective planet journalbearing 75. Rotary motion of the sun gear 68 urges each planet gear 70to rotate about a respective longitudinal axis P. The gears may begenerally as shown schematically in FIG. 1B.

Each planet gear 70 is also in meshed engagement with rotating ring gear74 that is mechanically connected to a fan shaft 76. Since the planetgears 70 mesh with both the rotating ring gear 74 as well as therotating sun gear 68, the planet gears 70 rotate about their own axes todrive the ring gear 74 to rotate about engine axis A. The rotation ofthe ring gear 74 is conveyed to the fan 42 (FIG. 1 ) through the fanshaft 76 to thereby drive the fan 42 at a lower speed than the low speedspool 30. It should be understood that the described geared architecture48 is but a single non-limiting embodiment and that various other gearedarchitectures will alternatively benefit herefrom.

With reference to FIG. 3 , a flexible support 78 supports the planetcarrier 72 to at least partially support the FDGS 60A with respect tothe static structure 36 such as a front center body which facilitatesthe segregation of vibrations and other transients therebetween. Itshould be understood that various gas turbine engine case structures mayalternatively or additionally provide the static structure and flexiblesupport 78. It is to be understood that the term “lateral” as usedherein refers to a perpendicular direction with respect to the axis ofrotation A and the term “transverse” refers to a pivotal bendingmovement with respect to the axis of rotation A so as to absorbdeflections which may be otherwise applied to the FDGS 60. The staticstructure 36 may further include a number 1 and 1.5 bearing supportstatic structure 82 which is commonly referred to as a “K-frame” whichsupports the number 1 and number 1.5 bearing systems 38A. 38B. Notably,the K-frame bearing support defines a lateral stiffness (represented asKframe in FIG. 3 ) and a transverse stiffness (represented as KframeBENDin FIG. 3 ) as the referenced factors in this non-limiting embodiment.

In this disclosed non-limiting embodiment, the lateral stiffness (KFS;KIC) of both the flexible support 78 and the input coupling 62 are eachless than about 11% of the lateral stiffness (Kframe). That is, thelateral stiffness of the entire FDGS 60 is controlled by this lateralstiffness relationship. Alternatively, or in addition to thisrelationship, the transverse stiffness of both the flexible support 78and the input coupling 62 are each less than about 11% of the transversestiffness (KframeBEND). That is, the transverse stiffness of the entireFDGS 60 is controlled by this transverse stiffness relationship.

With reference to FIG. 4 , another non-limiting embodiment of a FDGS 60Bincludes a flexible support 78′ that supports a rotationally fixed ringgear 74′. The fan shaft 76′ is driven by the planet carrier 72′ in theschematically illustrated planet system which otherwise generallyfollows the star system architecture of FIG. 3 .

With reference to FIG. 5 , the lateral stiffness relationship within aFDGS 60C itself (for a star system architecture) is schematicallyrepresented. The lateral stiffness (KIC) of an input coupling 62, alateral stiffness (KFS) of a flexible support 78, a lateral stiffness(KRG) of a ring gear 74 and a lateral stiffness (KJB) of a planetjournal bearing 75 are controlled with respect to a lateral stiffness(KGM) of a gear mesh within the FDGS 60.

In the disclosed non-limiting embodiment, the stiffness (KGM) may bedefined by the gear mesh between the sun gear 68 and the multiple planetgears 70. The lateral stiffness (KGM) within the FDGS 60 is thereferenced factor and the static structure 82′ rigidly supports the fanshaft 76. That is, the fan shaft 76 is supported upon bearing systems38A, 38B which are essentially rigidly supported by the static structure82′. The lateral stiffness (KJB) may be mechanically defined by, forexample, the stiffness within the planet journal bearing 75 and thelateral stiffness (KRG) of the ring gear 74 may be mechanically definedby, for example, the geometry of the ring gear wings 74L, 74R (FIG. 2 ).

In the disclosed non-limiting embodiment, the lateral stiffness (KRG) ofthe ring gear 74 is less than about 12% of the lateral stiffness (KGM)of the gear mesh; the lateral stiffness (KFS) of the flexible support 78is less than about 8% of the lateral stiffness (KGM) of the gear mesh;the lateral stiffness (KJB) of the planet journal bearing 75 is lessthan or equal to the lateral stiffness (KGM) of the gear mesh; and thelateral stiffness (KIC) of an input coupling 62 is less than about 5% ofthe lateral stiffness (KGM) of the gear mesh.

With reference to FIG. 6 , another non-limiting embodiment of a lateralstiffness relationship within a FDGS 60D itself are schematicallyillustrated for a planetary gear system architecture, which otherwisegenerally follows the star system architecture of FIG. 5 .

It should be understood that combinations of the above lateral stiffnessrelationships may be utilized as well. The lateral stiffness of each ofstructural components may be readily measured as compared to filmstiffness and spline stiffness which may be relatively difficult todetermine.

By flex mounting to accommodate misalignment of the shafts under designloads, the FDGS design loads have been reduced by more than 17% whichreduces overall engine weight. The flex mount facilitates alignment toincrease system life and reliability. The lateral flexibility in theflexible support and input coupling allows the FDGS to essentially‘float’ with the fan shaft during maneuvers. This allows: (a) the torquetransmissions in the fan shaft, the input coupling and the flexiblesupport to remain constant during maneuvers; (b) maneuver inducedlateral loads in the fan shaft (which may otherwise potentially misaligngears and damage teeth) to be mainly reacted to through the number 1 and1.5 bearing support K-frame; and (c) both the flexible support and theinput coupling to transmit small amounts of lateral loads into the FDGS.The splines, gear tooth stiffness, journal bearings, and ring gearligaments are specifically designed to minimize gear tooth stressvariations during maneuvers. The other connections to the FDGS areflexible mounts (turbine coupling, case flex mount). These mount springrates have been determined from analysis and proven in rig and flighttesting to isolate the gears from engine maneuver loads. In addition,the planet journal bearing spring rate may also be controlled to supportsystem flexibility.

FIG. 7 is similar to FIG. 5 but shows the transverse stiffnessrelationships within the FDGS 60C (for a star system architecture). Thetransverse stiffness (KICBEND) of the input coupling 62, a transversestiffness (KFSBEND) of the flexible support 78, a transverse stiffness(KRGBEND) of the ring gear 74 and a transverse stiffness (KJBBEND) ofthe planet journal bearing 75 are controlled with respect to atransverse stiffness (KGMBEND) of the gear mesh within the FDGS 60.

In the disclosed non-limiting embodiment, the stiffness (KGMBEND) may bedefined by the gear mesh between the sun gear 68 and the multiple planetgears 70. The transverse stiffness (KGMBEND) within the FDGS 60 is thereferenced factor and the static structure 82′ rigidly supports the fanshaft 76. That is, the fan shaft 76 is supported upon bearing systems38A, 38B which are essentially rigidly supported by the static structure82′. The transverse stiffness (KJBBEND) may be mechanically defined by,for example, the stiffness within the planet journal bearing 75 and thetransverse stiffness (KRGBEND) of the ring gear 74 may be mechanicallydefined by, for example, the geometry of the ring gear wings 74L, 74R(FIG. 2 ).

In the disclosed non-limiting embodiment, the transverse stiffness(KRGBEND) of the ring gear 74 is less than about 12% of the transversestiffness (KGMBEND) of the gear mesh; the transverse stiffness (KFSBEND)of the flexible support 78 is less than about 8% of the transversestiffness (KGMBEND) of the gear mesh; the transverse stiffness (KJBBEND)of the planet journal bearing 75 is less than or equal to the transversestiffness (KGMBEND) of the gear mesh; and the transverse stiffness(KICBEND) of an input coupling 62 is less than about 5% of thetransverse stiffness (KGMBEND) of the gear mesh.

FIG. 8 is similar to FIG. 6 but shows the transverse stiffnessrelationship within the FDGS 60D for the planetary gear systemarchitecture.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The combined arrangement of the high power density and fan drive turbinewith the high AN2 performance quantity, all incorporated with theflexible mounting structure, provide a very robust and efficient gasturbine engine.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

1. A turbofan engine comprising: a propulsor section including apropulsor shaft in driving engagement with a propulsor having bladesextending from a propulsor hub and being rotatable about a centrallongitudinal axis of said turbofan engine; an epicyclic gear system indriving engagement with said propulsor shaft and having a gear meshlateral stiffness and a gear mesh transverse stiffness; a gear systeminput to said epicyclic gear system defining a gear system input lateralstiffness and a gear system input transverse stiffness, wherein saidgear system input lateral stiffness is less than 5% of said gear meshlateral stiffness; a first turbine section in driving engagement withsaid gear system input and rotatable about said central longitudinalaxis and said first turbine section includes an inlet and an outlet; anda second turbine section, wherein said first turbine section has a firstexit area at a first exit point at said outlet and rotates at a firstspeed, wherein said second turbine section has a second exit area at asecond exit point and rotates at a second speed, which is faster thansaid first speed, said first and second speeds being redline speeds,wherein a first performance quantity is defined as the product of saidfirst speed squared and said first area, wherein a second performancequantity is defined as the product of said second speed squared and saidsecond area, wherein a performance quantity ratio of said firstperformance quantity to said second performance quantity is between 0.5and 1.5.
 2. The turbofan engine as set forth in claim 1, wherein saidepicyclic gear system includes a ring gear having a ring gear lateralstiffness and a ring gear transverse stiffness, and said ring gearlateral stiffness is less than 12% of said gear mesh lateral stiffness.3. The turbofan engine as set forth in claim 2, wherein said firstturbine has between three and six stages and said second turbine has twostages.
 4. The turbofan engine as set for in claim 3, wherein saidsecond speed is more than twice said first speed.
 5. The turbofan engineas set forth in claim 3, wherein there being a flexible support tosupport said epicyclic gear system, said flexible support having aflexible support lateral stiffness and a flexible support transversestiffness, a frame which supports said fan shaft having a frame lateralstiffness and a frame transverse stiffness, and at least one of saidflexible support lateral stiffness and said flexible support transversestiffness being less than 11% of a respective one of said frame lateralstiffness and said frame transverse stiffness.
 6. The turbofan engine asset forth in claim 5, wherein said gear system input connects a sun gearof said epicyclic gear system to be driven by said first turbinesection, and said gear system input lateral stiffness and said gearsystem input transverse stiffness being less than 11% of a respectiveone of said frame lateral stiffness and said frame transverse stiffness.7. The turbofan engine as set forth in claim 5, wherein said performancequantity ratio is greater than or equal to 0.5 and less than or equal to1.075.
 8. The turbofan engine as set forth in claim 5, wherein said gearsystem input transverse stiffness is less than 5% of said gear meshtransverse stiffness.
 9. The turbofan engine as set forth in claim 1,wherein said first turbine has between three and six stages and saidsecond turbine has two stages.
 10. The turbofan engine as set forth inclaim 9, wherein there being a flexible support to support saidepicyclic gear system, said flexible support having a flexible supportlateral stiffness and a flexible support transverse stiffness, and atleast one of said flexible support lateral stiffness and said flexiblesupport transverse stiffness being less than 8% of a respective one ofsaid gear mesh lateral stiffness and said gear mesh transversestiffness.
 11. The turbofan engine as set forth in claim 9, whereinthere being a flexible support to support said epicyclic gear system,said flexible support having a flexible support lateral stiffness and aflexible support transverse stiffness, and said flexible support lateralstiffness and said flexible support transverse stiffness being less than8% of a respective one of said gear mesh lateral stiffness and said gearmesh transverse stiffness.
 12. The turbofan engine as set for in claim1, wherein the propulsor section is a fan section, the propulsor is afan, the propulsor hub is a fan hub, and an outer housing surrounds thefan to define a bypass duct.
 13. A turbofan engine comprising: apropulsor section including a propulsor shaft in driving engagement witha propulsor having blades extending from a propulsor hub and beingrotatable about a central longitudinal axis of said turbofan engine; anepicyclic gear system in driving engagement with said propulsor shaftand having a gear mesh lateral stiffness and a gear mesh transversestiffness; and a gear system input to said epicyclic gear systemdefining a gear system input lateral stiffness and a gear system inputtransverse stiffness, wherein said gear system input transversestiffness is less than 5% of said gear mesh transverse stiffness; afirst turbine section in driving engagement with said gear system inputand rotatable about said central longitudinal axis and said firstturbine section includes an inlet and an outlet; and a second turbinesection, wherein said first turbine section has a first exit area at afirst exit point at said outlet and rotates at a first speed, whereinsaid second turbine section has a second exit area at a second exitpoint and rotates at a second speed, which is faster than said firstspeed, said first and second speeds being redline speeds, wherein afirst performance quantity is defined as the product of said first speedsquared and said first area, wherein a second performance quantity isdefined as the product of said second speed squared and said secondarea, wherein a performance quantity ratio of said first performancequantity to said second performance quantity is between 0.5 and 1.5. 14.The turbofan engine as set forth in claim 13, wherein said epicyclicgear system includes a ring gear having a ring gear lateral stiffnessand a ring gear transverse stiffness and said ring gear lateralstiffness is less than 12% of said gear mesh lateral stiffness.
 15. Theturbofan engine as set forth in claim 14, wherein there being a flexiblesupport to support said epicyclic gear system, said flexible supporthaving a flexible support lateral stiffness and a flexible supporttransverse stiffness, and at least one of said flexible support lateralstiffness and said flexible support transverse stiffness being less than8% of a respective one of said gear mesh lateral stiffness and said gearmesh transverse stiffness.
 16. The turbofan engine as set forth in claim15, wherein said first turbine has between three and six stages and saidsecond turbine has two stages.
 17. The turbofan engine as set forth inclaim 16, gear system input lateral stiffness is less than 5% of saidgear mesh lateral stiffness.
 18. The turbofan engine as set forth inclaim 17, wherein at least one of said flexible support lateralstiffness and said flexible support transverse stiffness is less than11% of a respective one of a frame lateral stiffness and a frametransverse stiffness of a frame which supports said fan shaft.
 19. Theturbofan engine as set for in claim 18, wherein said second speed ismore than twice said first speed.
 20. The turbofan engine as set forthin claim 14, wherein at least one of said flexible support lateralstiffness and said flexible support transverse stiffness is less than11% of a respective one of a frame lateral stiffness and a frametransverse stiffness of a frame which supports said fan shaft.
 21. Theturbofan engine as set forth in claim 20, wherein a frame which supportssaid fan shaft defines a frame lateral stiffness and a frame transversestiffness, and said gear system input lateral stiffness and said gearsystem input transverse stiffness are less than 11% of a respective oneof said frame lateral stiffness and said frame transverse stiffness. 22.The turbofan engine as set forth in claim 21, wherein said gear systeminput transverse stiffness is less than 5% of said gear mesh transversestiffness.
 23. The turbofan engine as set forth in claim 12, whereinsaid first turbine section having between three and six stages and saidsecond tur bine section having two stages.
 24. The turbofan engine asset forth in claim 23, wherein, said epicyclic gear system includes aring gear having a ring gear lateral stiffness and a ring geartransverse stiffness and said ring gear lateral stiffness is less than12% of said gear mesh lateral stiffness.
 25. The turbofan engine as setforth in claim 24, wherein said ring gear transverse stiffness is lessthan 12% of said gear mesh transverse stiffness.
 26. The turbofan engineas set forth in claim 24, wherein there being a flexible support tosupport said epicyclic gear system, said flexible support having aflexible support lateral stiffness and a flexible support transversestiffness, and at least one of said flexible support lateral stiffnessand said flexible support transverse stiffness being less than 8% of arespective one of said gear mesh lateral stiffness and said gear meshtransverse stiffness.
 27. The turbofan engine as set forth in claim 26,wherein at least one of said flexible support lateral stiffness and saidflexible support transverse stiffness is less than 11% of a respectiveone of a frame lateral stiffness and a frame transverse stiffness of aframe which supports said fan shaft.
 28. The turbofan engine as setforth in claim 26, wherein a frame which supports said fan shaft definesa frame lateral stiffness and a frame transverse stiffness and said gearsystem input lateral stiffness and said gear system input transversestiffness are less than 11% of a respective one of said frame lateralstiffness and said frame transverse stiffness.
 29. The turbofan engineas set forth in claim 28, wherein said performance quantity ratio isabove or equal to 0.8 and less than or equal to 1.075.
 30. The turbofanengine as set for in claim 29, wherein the propulsor section is a fansection, the propulsor is a fan, the propulsor hub is a fan hub, and anouter housing surrounds the fan to define a bypass duct.